In situ gas turbine prevention of crack growth progression

ABSTRACT

A method for remotely stopping a crack in a component of a gas turbine engine is provided. The method can include inserting an integrated repair interface attached to a cable delivery system within a gas turbine engine; positioning the tip adjacent to a defect within a surface of the component; temporarily attaching the tip adjacent to the defect within the surface on the component; supplying a new material to the area to fill the defect; and heating the new material to fuse the new material to the component within the defect.

FIELD OF THE INVENTION

The present subject matter relates generally to gas turbine engines and,more particularly, to a system and method for performing an in siturepair of an internal component of a gas turbine engine.

BACKGROUND OF THE INVENTION

A gas turbine engine typically includes a turbomachinery core having ahigh pressure compressor, combustor, and high pressure turbine in serialflow relationship. The core is operable in a known manner to generate aprimary gas flow. The high pressure compressor includes annular arrays(“rows”) of stationary vanes that direct air entering the engine intodownstream, rotating blades of the compressor. Collectively one row ofcompressor vanes and one row of compressor blades make up a “stage” ofthe compressor. Similarly, the high pressure turbine includes annularrows of stationary nozzle vanes that direct the gases exiting thecombustor into downstream, rotating blades of the turbine. Collectivelyone row of nozzle vanes and one row of turbine blades make up a “stage”of the turbine. Typically, both the compressor and turbine include aplurality of successive stages.

Gas turbine engines, particularly aircraft engines, require a highdegree of periodic maintenance. For example, periodic maintenance isoften scheduled to allow internal components of the engine to beinspected for defects and subsequently repaired. Unfortunately, manyconventional repair methods used for aircraft engines require that theengine be removed from the body of the aircraft and subsequentlypartially or fully disassembled. As such, these repair methods result ina significant increase in both the time and the costs associated withrepairing internal engine components.

Accordingly, a system and method for performing an in situ repair of aninternal component of a gas turbine engine would be welcomed within thetechnology.

BRIEF DESCRIPTION OF THE INVENTION

Aspects and advantages of the invention will be set forth in part in thefollowing description, or may be obvious from the description, or may belearned through practice of the invention.

A method is generally provided for remotely stopping a crack in acomponent of a gas turbine engine. In one embodiment, the methodincludes inserting an integrated repair interface attached to a cabledelivery system within a gas turbine engine; positioning the tipadjacent to a defect within a surface of the component; temporarilyattaching the tip adjacent to the defect within the surface on thecomponent; supplying a new material to the area to fill the defect; andheating the new material to fuse the new material to the componentwithin the defect.

These and other features, aspects and advantages of the presentinvention will become better understood with reference to the followingdescription and appended claims. The accompanying drawings, which areincorporated in and constitute a part of this specification, illustrateembodiments of the invention and, together with the description, serveto explain the principles of the invention.

BRIEF DESCRIPTION OF THE DRAWINGS

A full and enabling disclosure of the present invention, including thebest mode thereof, directed to one of ordinary skill in the art, is setforth in the specification, which makes reference to the appended Figs.,in which:

FIG. 1 illustrates a cross-sectional view of one embodiment of a gasturbine engine that may be utilized within an aircraft in accordancewith aspects of the present subject matter;

FIG. 2 illustrates a partial, cross-sectional view of one embodiment ofa turbine suitable for use within the gas turbine engine shown in FIG.1, particularly illustrating access ports defined in the engine forproviding internal access to the turbine;

FIG. 3 illustrates a partial, cross-sectional view of one embodiment ofa compressor suitable for use within the gas turbine engine shown inFIG. 1, particularly illustrating access ports defined in the engine forproviding internal access to the compressor;

FIG. 4 illustrates a simplified view of one embodiment of a system forperforming an in situ repair of an internal component of a gas turbineengine in accordance with aspects of the present subject matter,particularly illustrating a repair tool inserted through an access portof the engine to access a defect of the internal component;

FIG. 5 illustrates a partial view of an exemplary repair tooltemporarily secured to a tip of an airfoil in order to perform an insitu repair thereon;

FIG. 6 illustrates a partial view of another exemplary repair tooltemporarily secured to a tip of an airfoil in order to perform an insitu repair thereon;

FIG. 7 illustrates a partial view of one embodiment of the repair tooltemporarily secured to a surface of an internal component of the gasturbine engine in order to supply powder particles within a defect forin situ repair; and

FIG. 8 illustrates a simplified view of one embodiment of a system forperforming an in situ repair of an internal component of a gas turbineengine in accordance with aspects of the present subject matter,particularly illustrating a repair tool inserted through an access portof the engine to access a defect of the internal component and supply afill material into a defect on the component.

Repeat use of reference characters in the present specification anddrawings is intended to represent the same or analogous features orelements of the present invention.

DETAILED DESCRIPTION OF THE INVENTION

Reference now will be made in detail to embodiments of the invention,one or more examples of which are illustrated in the drawings. Eachexample is provided by way of explanation of the invention, notlimitation of the invention. In fact, it will be apparent to thoseskilled in the art that various modifications and variations can be madein the present invention without departing from the scope or spirit ofthe invention. For instance, features illustrated or described as partof one embodiment can be used with another embodiment to yield a stillfurther embodiment. Thus, it is intended that the present inventioncovers such modifications and variations as come within the scope of theappended claims and their equivalents.

As used herein, the terms “first”, “second”, and “third” may be usedinterchangeably to distinguish one component from another and are notintended to signify location or importance of the individual components.

The terms “upstream” and “downstream” refer to the relative directionwith respect to fluid flow in a fluid pathway. For example, “upstream”refers to the direction from which the fluid flows, and “downstream”refers to the direction to which the fluid flows.

In general, a system and method is provided for performing an in siturepair of an internal component of a gas turbine engine. In severalembodiments, the system may include a repair tool configured to beinserted through an access port of the gas turbine engine to allow arepair tip or tip end of the tool to be positioned adjacent to a defectof an internal component of the engine, such as a crack, void,distressed area or any other defect defining a fillable volume. As willbe described below, the repair tool may be configured to temporarilyattach to the surface of the component, allowing precision work to beperformed on the component. For example, the repair tool can supply anew material (solid or liquid) and/or a heating element to fill and fusenew material within the crack to repair the defect.

It should be appreciated that the disclosed system and method maygenerally be used to perform in situ repairs of internal componentslocated within any suitable type of gas turbine engine, includingaircraft-based turbine engines and land-based turbine engines,regardless of the engine's current assembly state (e.g., fully orpartially assembled). Additionally, with reference to aircraft engines,it should be appreciated that the present subject matter may beimplemented on-wing or off-wing.

Referring now to the drawings, FIG. 1 illustrates a cross-sectional viewof one embodiment of a gas turbine engine 10 that may be utilized withinan aircraft in accordance with aspects of the present subject matter,with the engine 10 being shown having a longitudinal or axial centerlineaxis 12 extending therethrough for reference purposes. In general, theengine 10 may include a core gas turbine engine (indicated generally byreference character 14) and a fan section 16 positioned upstreamthereof. The core engine 14 may generally include a substantiallytubular outer casing 18 that defines an annular inlet 20. In addition,the outer casing 18 may further enclose and support a booster compressor22 for increasing the pressure of the air that enters the core engine 14to a first pressure level. A high pressure, multi-stage, axial-flowcompressor 24 may then receive the pressurized air from the boostercompressor 22 and further increase the pressure of such air. Thepressurized air exiting the high-pressure compressor 24 may then flow toa combustor 26 within which fuel is injected into the flow ofpressurized air, with the resulting mixture being combusted within thecombustor 26. The high energy combustion products are directed from thecombustor 26 along the hot gas path of the engine 10 to a first (highpressure) turbine 28 for driving the high pressure compressor 24 via afirst (high pressure) drive shaft 30, and then to a second (lowpressure) turbine 32 for driving the booster compressor 22 and fansection 16 via a second (low pressure) drive shaft 34 that is generallycoaxial with first drive shaft 30. After driving each of turbines 28 and32, the combustion products may be expelled from the core engine 14 viaan exhaust nozzle 36 to provide propulsive jet thrust.

Additionally, as shown in FIG. 1, the fan section 16 of the engine 10may generally include a rotatable, axial-flow fan rotor assembly 38 thatis configured to be surrounded by an annular fan casing 40. It should beappreciated by those of ordinary skill in the art that the fan casing 40may be configured to be supported relative to the core engine 14 by aplurality of substantially radially-extending, circumferentially-spacedoutlet guide vanes 42. As such, the fan casing 40 may enclose the fanrotor assembly 38 and its corresponding fan rotor blades 44. Moreover, adownstream section 46 of the fan casing 40 may extend over an outerportion of the core engine 14 so as to define a secondary, or by-pass,airflow conduit 48 that provides additional propulsive jet thrust.

It should be appreciated that, in several embodiments, the second (lowpressure) drive shaft 34 may be directly coupled to the fan rotorassembly 38 to provide a direct-drive configuration. Alternatively, thesecond drive shaft 34 may be coupled to the fan rotor assembly 38 via aspeed reduction device 37 (e.g., a reduction gear or gearbox) to providean indirect-drive or geared drive configuration. Such a speed reductiondevice(s) may also be provided between any other suitable shafts and/orspools within the engine 10 as desired or required.

During operation of the engine 10, it should be appreciated that aninitial air flow (indicated by arrow 50) may enter the engine 10 throughan associated inlet 52 of the fan casing 40. The air flow 50 then passesthrough the fan blades 44 and splits into a first compressed air flow(indicated by arrow 54) that moves through conduit 48 and a secondcompressed air flow (indicated by arrow 56) which enters the boostercompressor 22. The pressure of the second compressed air flow 56 is thenincreased and enters the high pressure compressor 24 (as indicated byarrow 58). After mixing with fuel and being combusted within thecombustor 26, the combustion products 60 exit the combustor 26 and flowthrough the first turbine 28. Thereafter, the combustion products 60flow through the second turbine 32 and exit the exhaust nozzle 36 toprovide thrust for the engine 10.

The gas turbine engine 10 may also include a plurality of access portsdefined through its casings and/or frames for providing access to theinterior of the core engine 14. For instance, as shown in FIG. 1, theengine 10 may include a plurality of access ports 62 (only six of whichare shown) defined through the outer casing 18 for providing internalaccess to one or both of the compressors 22, 24 and/or for providinginternal access to one or both of the turbines 28, 32. In severalembodiments, the access ports 62 may be spaced apart axially along thecore engine 14. For instance, the access ports 62 may be spaced apartaxially along each compressor 22, 24 and/or each turbine 28, 32 suchthat at least one access port 62 is located at each compressor stageand/or each turbine stage for providing access to the internalcomponents located at such stage(s). In addition, the access ports 62may also be spaced apart circumferentially around the core engine 14.For instance, a plurality of access ports 62 may be spaced apartcircumferentially around each compressor stage and/or turbine stage.

It should be appreciated that, although the access ports 62 aregenerally described herein with reference to providing internal accessto one or both of the compressors 22, 24 and/or for providing internalaccess to one or both of the turbines 28, 32, the gas turbine engine 10may include access ports 62 providing access to any suitable internallocation of the engine 10, such as by including access ports 62 thatprovide access within the combustor 26 and/or any other suitablecomponent of the engine 10.

Referring now to FIG. 2, a partial, cross-sectional view of the first(or high pressure) turbine 28 described above with reference to FIG. 1is illustrated in accordance with embodiments of the present subjectmatter. As shown, the first turbine 28 may include a first stage turbinenozzle 66 and an annular array of rotating turbine blades 68 (one ofwhich is shown) located immediately downstream of the nozzle 66. Thenozzle 66 may generally be defined by an annular flow channel thatincludes a plurality of radially-extending, circularly-spaced nozzlevanes 70 (one of which is shown). The vanes 70 may be supported betweena number of arcuate outer bands 72 and arcuate inner bands 74.Additionally, the circumferentially spaced turbine blades 68 maygenerally be configured to extend radially outwardly from a rotor disk(not shown) that rotates about the centerline axis 12 (FIG. 1) of theengine 10. Moreover, a turbine shroud 76 may be positioned immediatelyadjacent to the radially outer tips of the turbine blades 68 so as todefine the outer radial flowpath boundary for the combustion products 60flowing through the turbine 28 along the hot gas path of the engine 10.

As indicated above, the turbine 28 may generally include any number ofturbine stages, with each stage including an annular array of nozzlevanes and follow-up turbine blades 68. For example, as shown in FIG. 2,an annular array of nozzle vanes 78 of a second stage of the turbine 28may be located immediately downstream of the turbine blades 68 of thefirst stage of the turbine 28.

Moreover, as shown in FIG. 2, a plurality of access ports 62 may bedefined through the turbine casing and/or frame, with each access port62 being configured to provide access to the interior of the turbine 28at a different axial location. Specifically, as indicated above, theaccess ports 62 may, in several embodiments, be spaced apart axiallysuch that each access port 62 is aligned with or otherwise providesinterior access to a different stage of the turbine 28. For instance, asshown in FIG. 2, a first access port 62A may be defined through theturbine casing/frame to provide access to the first stage of the turbine28 while a second access port 62B may be defined through the turbinecasing/frame to provide access to the second stage of the turbine 28.

It should be appreciated that similar access ports 62 may also beprovided for any other stages of the turbine 28 and/or for any turbinestages of the second (or low pressure) turbine 32. It should also beappreciated that, in addition to the axially spaced access ports 62shown in FIG. 2, access ports 62 may be also provided at differingcircumferentially spaced locations. For instance, in one embodiment, aplurality of circumferentially spaced access ports may be definedthrough the turbine casing/frame at each turbine stage to provideinterior access to the turbine 28 at multiple circumferential locationsaround the turbine stage.

Referring now to FIG. 3, a partial, cross-sectional view of the highpressure compressor 24 described above with reference to FIG. 1 isillustrated in accordance with embodiments of the present subjectmatter. As shown, the compressor 24 may include a plurality ofcompressor stages, with each stage including both an annular array offixed compressor vanes 80 (only one of which is shown for each stage)and an annular array of rotatable compressor blades 82 (only one ofwhich is shown for each stage). Each row of compressor vanes 80 isgenerally configured to direct air flowing through the compressor 24 tothe row of compressor blades 82 immediately downstream thereof.

Moreover, the compressor 24 may include a plurality of access ports 62defined through the compressor casing/frame, with each access port 62being configured to provide access to the interior of the compressor 24at a different axial location. Specifically, in several embodiments, theaccess ports 62 may be spaced apart axially such that each access port62 is aligned with or otherwise provides interior access to a differentstage of the compressor 24. For instance, as shown in FIG. 3, first,second, third and fourth access ports 62 a, 62 b, 62 c, 62 d areillustrated that provide access to four successive stages, respectively,of the compressor 24.

It should be appreciated that similar access ports 62 may also beprovided for any of the other stages of the compressor 24 and/or for anyof the stages of the low pressure compressor 22. It should also beappreciated that, in addition to the axially spaced access ports 62shown in FIG. 3, access ports 62 may be also provided at differingcircumferentially spaced locations. For instance, in one embodiment, aplurality of circumferentially spaced access ports may be definedthrough the compressor casing/frame at each compressor stage to provideinterior access to the compressor 24 at multiple circumferentiallocations around the compressor stage.

Referring now to FIG. 4, a simplified view of one embodiment of a system100 for performing an in situ repair of an internal component of a gasturbine engine 10 are illustrated in accordance with aspects of thepresent subject matter. As shown, the system 100 may include a repairtool 102 configured to be inserted through an access port 62 of the gasturbine engine 10, such as any of the access ports 62 described abovewith reference to FIGS. 1-3, to allow an in situ repair procedure to beperformed on an internal component(s) (indicated by dashed lines 104) ofthe engine 10.

In general, the repair tool 102 may correspond to any suitable tool(s)and/or component(s) that may be inserted through an access port 62 ofthe gas turbine engine 10 and attach onto the surface 105 of thecomponent 104 to perform precision work thereon. For example, anattachment mechanism 135 can temporarily attach onto the surface 105 sothat the tool 102 can perform work at or near an identified defect 106of the internal engine component(s) 104 being repaired (e.g., a turbineblade(s)). As such, the repair tool 102 may be temporarily attached tothe surface 105 so as to allow for precision work at the defect 106(e.g., with precision accuracy within about 0.5 mm or less, such asabout 0.25 mm or less). As generically shown in FIG. 4, a conduit 110 isattached to a working head 122 includes a work mechanism 124controllable via a controller 114 (e.g., a computer or otherprogrammable machine).

In one embodiment, the attachment mechanism 135 can be a tripod grip fora component 104 having a known shape and/or size. As shown in FIG. 5,the component 104 is an airfoil tip 200 with a known shape and size(e.g., a nozzle and/or blade). In other embodiments, the component 104can be a trailing edge and/or leading edge of the airfoil. Theattachment mechanism 135 includes a plurality of grip arms 150 thatattach the repair tool 102 onto the surface 105. The grip arms 150 arebrought together onto the edge of the tip 200 until the repair tool 102is secured onto the tip 200. In the embodiment shown, three grip arms150 are included in the attachment mechanism 135, although any suitablenumber of grip arms 150 may be utilized (e.g., three or more grip arms).

In another embodiment, the attachment mechanism 135 can be a suction cupattached onto the repair tool 102. As shown in FIG. 6, the attachmentmechanism 135 includes a suction cup 160 that attach the repair tool 102onto the surface 105. In one embodiment, a vacuum can be applied withinthe suction cup 160 to hold the repair tool 102 onto the surface inplace. The suction cup 160 can be constructed of a deformable,air-impervious material (e.g., a rubber material) that can form asuction attachment with the surface 105. Although shown with one suctioncup 160, any number of suction cups can be utilized to secure the repairtool 102 onto the surface 105. In yet another embodiment, an adhesivecan be utilized to secure the repair tool 102 onto the surface 105, suchas a hot melt adhesive, epoxy material, etc. Then, the adhesive materialcan be melted to remove the repair tool 102 from the surface 105.

Through the attachment mechanism 135, the location of repair tool 102can be precisely controlled and temporarily secured in place, whichallows for precision work to be performed. In one embodiment, a workinghead 122 is positioned and secured adjacent to he identified defect 106of the internal engine component(s) 104 being repaired (e.g., a turbineblade(s)). For example, as particularly shown in FIG. 4, the defect 106corresponds to a crack, void or other defective area formed along theexterior of the component 104 that defines an open or fillable volume108 with a base 107 of the crack, void or other defective area.

As shown in FIGS. 5-9, the working head 122 includes a work mechanism124 configured for addressing the defect 106. In one embodiment, the newmaterial can be supplied from a location exterior to the engine to theinternal location of the defect 106 to allow the fillable volume 108defined by the defect 106 to be filled with the new material. FIG. 7shows the repair tool 102 configured to supply high velocity powderparticles 125 from the exterior of the engine into the fillable volume108 of the defect 106. Upon impacting a surface of the defect 106, thehigh velocity particles 125 may plastically deform and adhere to thesurface, thereby filling-in the fillable volume 108 and repairing thedefect 106. For example, the particles can impact the surface within thedefect 106 at a speed of about 150 meters per second (m/s) to about 900m/s.

The average size of the powder particles 125 can vary depending on theircomposition, gun type, nozzle type, gases used, etc. In mostembodiments, the particle size and distribution can be about 25 μm toabout 150 μm (e.g., about 35 μm to about 75 μm (i.e., 400 to about 200mesh)). In certain embodiments, no more than about five percent of theparticles are larger than about 75 μm (200 mesh) and no more than aboutfifteen percent of the particles being smaller than about 35 μm (400mesh).

The powder particles 125 can be supplied to the location of the defectvia the repair tool 102 such that the fillable volume 108 may befilled-in with the powder particles 125, thereby repairing the defect106. In several embodiments, the repair tool 102 may be configured tosupply the powder particles 125 within the interior of the gas turbineengine 10. For example, the powder particles 125 may be transported viathe repair tool 102 from a location exterior to the gas turbine engine10 to a location within the engine 10 to allow the powder particles 125to be injected or otherwise directed into the fillable volume 108defined by the defect 106.

The particles 125 may be supplied via a carrier fluid (e.g., a carriergas) that is inert to the coating deposition.

The powder particles 125 may then be heated to fuse the material withinthe fillable volume 108 to repair the defect 106. For example, therepair tool 102 may include a heating element at its working end to heatthe powder particles 125 prior to adhesion of the surface, therebyfilling in the fillable volume 108 to bond the material within thedefect 106. For example, the working head 122 may include a heatingcomponent to locally heat the base of the defect 106, before, during,and/or after deposition of the new material (e.g., the powder particles125). For example, the heating component may direct thermal energy intothe defect 106 in the surface 105 of the component 104. The heatingcomponent can heat a precision weld within the base 107 of the defect106 (e.g., at the deepest point from the surface 105 within thecomponent 104) to effectively stop the propagation of the defect 106through the component 104.

For example, the base 107 may be heated to a temperature of about 1000°C. to about 2000° C. (e.g., about 1800° C. to about 2000° C.),particularly with the component 104 is constructed from a metal alloy orsuper-alloy such as a nickel-based alloy, a chromium-based alloy, etc.

In one embodiment, the repair tool 102 may include one or more heatingelements (indicated by dashed lines 120) provided in operativeassociation within the high temperature conduit 110. As shown in theillustrated embodiment of FIG. 8, the repair tool 102 may include a hightemperature conduit 110 for transporting the metal particles fromoutside the engine 10 to the location of the defect 106. Specifically,as shown in FIG. 8, the high temperature conduit 110 may extendlengthwise between working head 122 located within the gas turbineengine 10 and a material supply end 114 located exterior to the engine10. The tip end of the tool 102 may generally be positioned adjacent tothe location of the defect 106 for directing the particles 125 into thefillable volume 108. Additionally, the material supply end 114 of thetool 102 may generally be configured to receive particles 125 from aparticle source. For example, as shown in FIG. 8, particles 125contained within a chamber (or other suitable powder particle source)located exterior to the gas turbine engine 10 may be supplied to thematerial supply end 114 of the tool 102. The particles 125 received atthe material supply end 114 may then be directed through the hightemperature conduit 110 to the tip end of the tool 102 to allow themetal particles to be delivered to the location of the defect 106.

It should be appreciated that the high temperature conduit 110 maygenerally be formed from any suitable high temperature material thatallows the conduit 110 to serve as a fluid delivery means for the liquidmetal. For example, in several embodiments, the high temperature conduit110 may be formed from a ceramic material capable of withstandingtemperatures above the melting temperature of the metal being suppliedto the defect 106. However, in other embodiments, the conduit 110 may beformed from any other suitable high temperature material.

In general, the heating element(s) 120 may be configured to generateheat within the high temperature conduit 110 as powder particles 125 isbeing supplied through the conduit 110 so as to allow for particle flowat the desired rate and speed. For example, in one embodiment, theheating element(s) 120 may correspond to a resisting heating element(s),such as one or more resistance wires, that is integrated into orincorporated within a wall(s) of the conduit 110. However, in anotherembodiment, the heating element(s) 120 may correspond to any othersuitable heat generating device(s) and/or component(s) that may be usedto provide heating within the conduit 110 so as to maintain thetemperature of the powder particles 125 at its desired deliverytemperature. In one embodiment, the particles 125 are delivered to thedefect 106 at a temperature within 25% of its melting point (e.g.,within 10% of its melting point).

It should be appreciated that the powder particles 125 may be composedof any suitable metal material. For example, in one embodiment, thepowder particles 125 may correspond to the parent metal material of theinternal component 104 being repaired. In other embodiments, the powderparticles 125 may correspond to any other metal material that issuitable for use as a repair material within a gas turbine engine 10.

In one embodiment, the repair tool 102 includes an optical probe 130adjacent to the working head 122 and configured to be used inassociation with the repair tool 102. For instance, as shown in FIG. 4,the optical probe 130 corresponds to a separate component configured tobe used in combination with the repair tool 102 for repairing the defect106. However, in other embodiments, the optical probe 130 may be coupledto or integrated within the repair tool 102. Additionally, as shown inFIG. 4, the optical probe 130 has been inserted through the same accessport 62 as the repair tool 102. However, in other embodiments, the probe130 may be inserted into a different access port 62 than the repair tool102, such as an access port 62 located adjacent to the access port 62within which the repair tool 102 has been inserted.

In general, the optical probe 130 may correspond to any suitable opticaldevice that allows images of the interior of the engine 10 to becaptured or otherwise obtained. For instance, in several embodiments,the optical probe 130 may correspond to a borescope, videoscope,fiberscope or any other similar optical device known in the art thatallows for the interior of a gas turbine engine 10 to be viewed throughan access port 62. In such embodiments, the optical probe 130 mayinclude one or more optical elements (indicated schematically by dashedbox 132), such as one or more optical lenses, optical fibers, imagecapture devices, cables, and/or the like, for obtaining views or imagesof the interior of the engine 10 at a tip 134 of the probe 130 and fortransmitting or relaying such images from the probe tip 134 along thelength of the probe 130 to the exterior of the engine 10 for viewing bythe personnel performing the repair procedure on the internalcomponent(s) 104. In addition, the probe 130 may include a light source(indicated by dashed box 136) positioned at or adjacent to the probe tip134 to provide lighting within the interior of the engine 10.

As shown in FIGS. 4 and 11, the optical probe 130 may also include anarticulation assembly 138 that allows the orientation of the probe tip134 to be adjusted within the interior of the gas turbine engine 10. Forexample, the articulation assembly 138 may allow for the probe tip 134to be rotated or pivoted about a single axis or multiple axes to adjustthe orientation of the tip 134 relative to the remainder of the probe130. It should be appreciated that the articulation assembly 138 maygenerally have any suitable configuration and/or may include anysuitable components that allow for adjustment of the orientation of theprobe tip 134 relative to the remainder of the probe 130. For example,in one embodiment, a plurality of articulation cables 140 may be coupledbetween the probe tip 134 and one or more articulation motors 142. Insuch an embodiment, by adjusting the tension of the cables 140 via themotor(s) 142, the probe tip 134 may be reoriented within the gas turbineengine 10.

In one particular embodiment, the articulation assembly 138 alsocontrols the attachment mechanism 135 so as to temporarily attach to thesurface 105 the component 104 in order to perform the desired workthereon.

Methods are generally provided for performing an in situ repair of aninternal component of a gas turbine engine. In general, the methods arediscussed herein with reference to the gas turbine engine 10 and thesystem 100 described above with reference to FIGS. 1-9. However, itshould be appreciated by those of ordinary skill in the art that thedisclosed methods may generally be implemented with gas turbine engineshaving any other suitable engine configuration and/or with systemshaving any other suitable system configuration. In addition, althoughthe methods are discussed in a particular order for purposes ofdiscussion, the methods discussed herein are not limited to anyparticular order or arrangement. One skilled in the art, using thedisclosures provided herein, will appreciate that various steps of themethods disclosed herein can be omitted, rearranged, combined, and/oradapted in various ways without deviating from the scope of the presentdisclosure.

The method may include inserting a repair tool through an access port ofthe gas turbine engine such that the tool includes a tip end positionedwithin the engine; positioning the tip adjacent to a defect (e.g., acrack or other distress point) within the surface of the component; andtemporarily attaching the tip adjacent to the defect to allow precisionwork to be performed. For example, as indicated above, the method mayinclude positioning the tip end of the repair tool adjacent to a defectof an internal component of the gas turbine engine. As indicated above,the defect 106 may, for example, correspond to a crack, void or otherdefective area of an internal component 104 of the gas turbine engine10.

Moreover, the method may include performing precision repair work (e.g.,supplying powder particles, heating, etc.) using the repair tool bytemporarily attaching the tip end of the repair tool to the surface ofthe component.

This written description uses examples to disclose the invention,including the best mode, and also to enable any person skilled in theart to practice the invention, including making and using any devices orsystems and performing any incorporated methods. The patentable scope ofthe invention is defined by the claims, and may include other examplesthat occur to those skilled in the art. Such other examples are intendedto be within the scope of the claims if they include structural elementsthat do not differ from the literal language of the claims, or if theyinclude equivalent structural elements with insubstantial differencesfrom the literal languages of the claims.

What is claimed is:
 1. A method of remotely stopping a crack in acomponent of a gas turbine engine, the method comprising: inserting anintegrated repair interface attached to a cable delivery system within agas turbine engine; positioning a tip adjacent to a defect within asurface of the component; temporarily attaching the tip adjacent to thedefect within the surface on the component; heating a high temperatureconduit with a heating element; supplying a new material to the hightemperature conduit from external of the engine, wherein the newmaterial is a particle powder comprising a plurality of solid particles;maintaining the new material in the high temperature conduit at adelivery temperature within about 25% of the melting point of the newmaterial; supplying the new material at the delivery temperature to aworking head; supplying the new material to the area to fill the defect;and heating the a base of the defect to fuse the new material to thecomponent within the defect.
 2. The method of claim 1, wherein theplurality of solid particles are supplied into the defect at a velocityof about 150 m/s to about 900 m/s.
 3. The method of claim 1, wherein theplurality of solid particles is supplied into the defect at an averageparticle size of about 25 μm to about 150 μm.
 4. The method of claim 1,wherein the plurality of solid particles is supplied into the defect atan average particle size of about 35 μm to about 75 μm.
 5. The method ofclaim 1, wherein the new material is supplied into the defect as aparticle powder at the delivery temperature.
 6. The method of claim 5,wherein the particle powder is supplied into the defect at a temperaturewithin about 10% of its melting point.
 7. The method of claim 1, whereinthe new material is supplied into the defect as a liquid.
 8. The methodof claim 1, wherein the high temperature conduit is formed from aceramic material.
 9. The method of claim 1, wherein the high temperatureconduit further comprises: a heating element integrated into a wall ofthe high temperature conduit.
 10. The method of claim 9, wherein theheating element is a resisting heating element comprising: a resistancewire.
 11. The method of claim 10 wherein the resisting heating elementprovides heating within the high temperature conduit so as to maintainthe delivery temperature.
 12. The method of claim 1, further comprising:inserting an optical probe through the access port or an adjacent accessport of the gas turbine engine, the optical probe being configured toprovide a view of the defect within the gas turbine engine.
 13. Themethod of claim 12, wherein positioning the tip end of the repair tooladjacent to the defect comprises positioning the tip end relative to thedefect based on the view provided by the optical probe.
 14. The methodof claim 1, wherein the tip is attached using a tripod grip, a glue, ora vacuum sucker.
 15. The method of claim 1 wherein the heating elementprovides heating within the high temperature conduit so as to maintainthe delivery temperature.
 16. The method of claim 1 wherein the hightemperature conduit comprises: a material having a melting point greaterthan the melting point of the new material.
 17. The method of claim 1,wherein the base of the defect is heated to a temperature of 1000° C. to2000° C.
 18. The method of claim 1 wherein the base of the defect isheated to a temperature of 1800° C. to 2000° C.